1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine blade with an internal cooling air circuit.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In an industrial gas turbine engine, a hot gas flow is passed through a turbine to produce mechanical work used to drive an electric generator for power production. The turbine generally includes four stages of stator vanes and rotor blades to convert the energy from the hot gas flow into mechanical energy that drives the rotor shaft of the engine. It is well known in the art of gas turbine engines that the efficiency of the engine can be increased by increasing the gas flow temperature that enters the turbine. However, the turbine inlet temperature is limited to the material properties and cooling capabilities of the turbine parts. This is especially important for the first stage turbine vanes and blades since these airfoils are exposed to the hottest gas flow.
In order to allow for higher temperatures, turbine blade designers have proposed several complex internal blade cooling circuits to maximize the blade cooling through the use of convection cooling, impingement cooling and film cooling of the blade. FIGS. 1 and 2 show a prior art turbine blade with an aft flowing triple pass all convection cooled cooling circuit design. The blade cooling circuit includes a first pass cooling channel located along the leading edge, a second pass in the blade mid-chord region, and a third pass channel located along the trailing edge. The third leg or pass of the serpentine cooling circuit includes pin fins extending across the walls of the blade to promote heat transfer, and includes a row of exit cooling holes or slots to discharge the cooling air from the serpentine flow circuit and out from the blade. In the FIG. 1 prior art turbine blade cooling circuit, the blade leading edge is cooled with the first up pass of the multi-pass channel flow, a direct feed leading edge. The leading edge cooling passage in general has a rough triangular cross sectional shape in the spanwise direction of the blade as seen in FIG. 2. The inner surface area of the leading edge cooling passage reduces in the cross sectional area to the apex of an acute angle. The distribution of the cooling flow to the leading edge corner decreases and the substantial flow velocity as well as the internal heat transfer coefficient is comparatively reduced.
Several prior art turbine blades show triple pass aft flowing serpentine cooling circuits such as U.S. Pat. No. 4,775,296 issued to Schwarzmann et al on Oct. 4, 1998 and entitled COOLABLE AIRFOIL FOR A ROTARY MACHINE; U.S. Pat. No. 4,515,526 issued to Levengood on May 7, 1985 and entitled COOLABLE AIRFOIL FOR A ROTARY MACHINE; U.S. Pat. No. 4,786,233 issued to Shizuya et al on Nov. 22, 1988 and entitled GAS TURBINE COOLED BLADE; U.S. Pat. No. 4,236,870 issued to Hucul, Jr. et al on Dec. 2, 1980 and entitled TURBINE BLADE; U.S. Pat. No. 4,416,585 issued to Abdel-Messeh on Nov. 22, 1983 and entitled BLADE COOLING FOR GAS TURBINE ENGINE; U.S. Pat. No. 4,278,400 issued to Yamarik et al on Jul. 14, 1981 and entitled COOLABLE ROTOR BLADE; U.S. Pat. No. 6,164,913 issued to Reddy on Dec. 26, 2000 and entitled DUST RESISTANT AIRFOIL COOLING; U.S. Pat. No. 5,503,527 issued to Lee et al on Apr. 2, 1996 and entitled TURBINE BLADE HAVING TIP SLOT; U.S. Pat. No. 4,589,824 issued to Kozlin on May 20, 1986 and entitled ROTOR BLADE HAVING A TIP CAP END CLOSURE; U.S. Pat. No. 5,403,157 issued to Moore on Apr. 4, 1995 and entitled HEAT EXCHANGE MEANS FOR OBTAINING TEMPERATURE GRADIENT BALANCE; and, U.S. Pat. No. 7,198,468 B2 issued to Papple on Apr. 3, 2007 and entitled INTERNALLY COOLED TURBINE BLADE. none of these prior art turbine blades address the issue of blade tip cooling and flow separation described below and in which the present invention addresses.
It is therefore an object of the present invention to provide for a new serpentine flow cooling circuit for a turbine blade mid-chord tip region that can be used in a first or second stage turbine blade.
It is another object of the present invention to provide for a first or a second stage turbine blade with a triple pass cooling circuit having a low tapered airfoil or wide open tip geometry.
It is another object of the present invention to provide for a turbine blade with an improved blade outer tip region cooling capability.